orbitalElements

Orbital elements of satellites in scenario

Syntax

elements = orbitalElements(sat)

Description

elements = orbitalElements(sat) returns the orbital elements of the specified satellite sat.

Input Arguments

sat — Satellite

row vector of Satellite objects

Satellite, specified as a row vector of Satellite objects.

Output Arguments

elements — Orbital elements

structure

Orbital elements of input sat, returned as a structure. The fields of the structure depend on the orbit propagator chosen using the OrbitPropagator property of the satelliteScenario object.

Two Body Keplerian

The two-body-keplerian orbit propagator has these fields:

• SemiMajorAxis

• Eccentricity

• Inclination

• RightAscensionOfAscendingNode

• ArgumentOfPeriapsis

• TrueAnomaly

• Period

SGP4 and SDP4

The sgp4 and sdp4 orbit propagators have these fields:

• Eccentricity

• Inclination

• RightAscensionOfAscendingNode

• ArgumentOfPeriapsis

• MeanAnomaly

• MeanMotion

• Epoch

• BStar

• Period

The orbital elements represent the mean values at Epoch.

Ephemeris

The ephemeris propagator has these fields:

• EphemerisStartTime

• EphemerisStopTime

• PositionTimeTable

• VelocityTimeTable